FIG. 1 shows a ducted fan gas turbine engine 10 comprising, in axial flow series: an air intake 12, a propulsive fan 14 having a plurality of fan blades 16, an intermediate pressure compressor 18, a high-pressure compressor 20, a combustor 22, a high-pressure turbine 24, an intermediate pressure turbine 26, a low-pressure turbine 28 and a core exhaust nozzle 30. A nacelle (not shown) generally surrounds a fan casing 32 and engine 10 and defines the intake 12, a bypass duct 34 and a bypass exhaust nozzle. The engine has a principal axis of rotation 31.
Air entering the intake 12 is accelerated by the fan 14 to produce a bypass flow and a core flow. The bypass flow travels down the bypass duct 34 and exits the bypass exhaust nozzle 36 to provide the majority of the propulsive thrust produced by the engine 10. The core flow enters in axial flow series the intermediate pressure compressor 18, high pressure compressor 20 and the combustor 22, where fuel is added to the compressed air and the mixture burnt. The hot combustion products expand through and drive the high, intermediate and low-pressure turbines 24, 26, 28 before being exhausted through the nozzle 30 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines 24, 26, 28 respectively drive the high and intermediate pressure compressors 20, 18 and the fan 14 by interconnecting shafts 38, 40, 42.
As is well known in the art, the turbines and compressors are each constructed from linear cascades of stator-rotors pairs. The stators and rotors are provided in flow series pairs such that the stator vanes align the upstream air flow to an optimum angle for interaction with the rotor blades. FIG. 2 shows a partial longitudinal section of a turbine interior 210 from an engine similar to that shown in FIG. 1. The turbine section shows a portion of a turbine blade 212 mounted on a rotor disc 214 which is rotatable about the principal axis of the engine 31. The flow direction is left to right as shown in the Figure such that the aero foil portion 216 of the blade receives the air from upstream stator vane (not shown) which is located to the left of the blade 212. The flow of hot air drives the blade 212 and disc 214 into a rotation and provides an axial rearward force. Aft of the rotor is a stator assembly 218. The stator assembly 218 includes the stator vane 219 of the next turbine stage. The stator vane includes an aerofoil portion (not shown) and platform 220 as is known in the art.
A partitioning wall 222 extends from the stator structure radially inwards towards the centre of the engine. The partitioning wall 222 includes a stationary part 224 and a rotating part 226 which are divided by a seal arrangement 228 to allow the required relative rotation. The partitioning wall 222 provides a separation between a first chamber and a second chamber 232. The two chambers are held at respective first and second operational pressures when in use. The first pressure of the first, upstream chamber is higher than the second pressure of the second, downstream chamber 232. The operating pressures are provided by compressor air which is bled from respective stages of the compressor.
The first chamber is constructed from a first wall in the form of the rotor disc 214 and a second wall which is the partition wall 222. The second chamber 232 is constructed from the partition wall 222 and a rotor disc 214 of a second downstream rotor (not shown). The radially outboard end of the first chamber is provided with a gap 234 which separates the rotor from the stator structure to allow for the relative rotation.
The gap 234 includes a seal member in the form of a swan neck rim seal provided by an annular flange which extends axially from the stator assembly towards the rotor. The seal member is located with a suitable clearance which provides a minimal operational separation between the rotor and the stator, thus allowing reasonably sealed and unhindered rotation.
The rotor forms part of a larger rotor structure (not shown). The attachment between the rotor and the larger rotor structure is via a short shaft and flanged connection 238. The flanged connection is fixedly mounted to a corresponding structure with any suitable coupling structure such as an array of bolts.
In operation, hot gas expelled from the combustor expands through the turbine and drives the rotor round and loads it rearwards. The first chamber is provided with reasonably high pressure cooling air bled from the compressor. The pressure of the cooling air is higher than the main gas flow path such that there is a positive pressure head within the engine core and prevents egress of hot gas outside of the protective environment of the main gas path annulus.
The second chamber 232 is provided with cooling air but of a lower pressure relative to the first chamber to reflect the local pressure of the main gas path at that location.
In some failure modes of the engine, for example a shaft break or disconnection of join 238, the rearward loading of the rotor causes it to move backwards and engage with the stator structure 218. In particular, the seal member clashes with the rotor, and trailing edge of the blade platform 240 and leading edge of the stator vane platform 220 collide and mesh together to bring the rotor to a desirable controlled and rapid halt. However, the contact between the seal member and rotor, seals the first chamber creating a pneumatic buffer which dampens the rearward movement of the rotor and binding of the rotating and static structures. This reduces the deceleration of the rotor which is undesirable.
The present invention seeks to provide an improved turbine arrangement which collapses more readily in a failure mode.